This application relates generally to gas turbine engine rotor blades and, more particularly, to methods and apparatus for fabricating a rotor assemblies.
Known gas turbine engine compressor rotor blades include airfoils having a leading edge, a trailing edge, a pressure side, a suction side, a root portion, and a tip portion. The pressure and suction sides connect at the airfoil leading and trailing edges, and span radially between the root and tip portions. An inner flow-path is defined at least partially by the root portion, and an outer flow-path is defined at least partially by a stationary casing coupled radially outward from the rotor blades. At least some known stationary casings include an abradable material that is spaced circumferentially within the casing and radially outward from the blade tip portion. At least some known compressors, for example, include a plurality of rows of rotor blades that extend radially and orthogonally outward from a rotor disk.
At least some known compressor rotor blades are coupled in a converging flow-path that may be susceptible to high airfoil radial loading and vibratory stresses generated by blade dynamic responses if the airfoil tips rub against the abradable casing. More specifically, such loading and stresses may be generated as a result of the rotor blade deflecting and rubbing the abradable casing. The blade dynamic response generally causes the airfoils to assume a first flex mode shape which results in high airfoil stresses at a peak location near the root portion of the airfoil. Moreover, generally the effect of tip rubs may be more severe to the airfoil when the suction side contacts the abradable casing rather than the pressure side.